This invention relates to health monitoring of gas turbine engines.
Conventional gas turbine engines use a fuel pump to deliver fuel to their combustion systems. In certain circumstances the fuel pump is susceptible to cavitation erosion during normal service operation. In some cases this can lead to loss of fuel pump performance and subsequent service disruption to the customer.
Currently, during engine shutdown the Engine Electronic Controller (EEC) records and sends a warning message when insufficient fuel pressure is available to drive the Hydro Mechanical Unit (HMU) dump valve. However, the prime function of this EEC message is to monitor the operation of the HMU dump valve and not to indicate poor fuel pump performance.
This conventional method only indirectly identifies loss of fuel pump performance i.e. when the operation of the HMU dump valve is compromised. Thus a major disadvantage of this conventional method is that there is no warning of imminent failure of the fuel pump. This has a serious implication in that an aircraft mission may be aborted due to failure of the fuel pump. Furthermore, this will necessitate an immediate and time-consuming replacement fuel pump and delays.
Further to this, the conventional method is not output as a parameter within the aircraft/engine health monitoring system. This does not allow monitoring of the reduction in HP fuel pump performance to be carried out prior to a failure warning message being set.
A further disadvantage is that a fleet of aircraft/engines does not allow fleet management and prioritised unit repair/removal to be carried out.